CN107782495A - Supersonic speed exerciser is used in a kind of static probe calibration - Google Patents
Supersonic speed exerciser is used in a kind of static probe calibration Download PDFInfo
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- G01L—MEASURING FORCE, STRESS, TORQUE, WORK, MECHANICAL POWER, MECHANICAL EFFICIENCY, OR FLUID PRESSURE
- G01L27/00—Testing or calibrating of apparatus for measuring fluid pressure
- G01L27/002—Calibrating, i.e. establishing true relation between transducer output value and value to be measured, zeroing, linearising or span error determination
- G01L27/005—Apparatus for calibrating pressure sensors
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Abstract
Supersonic speed exerciser is used in a kind of static probe calibration disclosed by the invention, belongs to supersonic speed static probe calibration field.The present invention includes supersonic nozzle, test section, exhaust section and standard static probe；Test section air inlet oral area is adapted with supersonic nozzle gas outlet, and test section gas outlet is adapted with exhaust section air inlet；Supersonic nozzle and test section fixed seal connection, test section and exhaust section fixed seal connection, standard static probe and the symmetrical holes for being inserted both sides above and below test section respectively by school static probe, probe and test section fixing seal at patchhole.Invention additionally discloses supersonic nozzle preferred implementation method.The present invention provides a kind of exerciser for carrying out calibration to static probe under supersonic airstream environment, for obtaining the static pressure coefficient of static probe, insensitivity of the probe to flow-deviation angle degree is tested, for improving the confidence level and measurement accuracy of test data in actual use.
Description
Technical field
The invention belongs to supersonic speed static probe to calibrate field, is related to a kind of calibrating installation for being used to calibrate static probe,
More particularly to supersonic speed exerciser is used in a kind of static probe calibration.
Background technology
Static probe used by measurement static pressure, error can be all produced in design, processing, assembling process, so that
Surveyed static pressure result deviates actual value.In order to ensure static pressure measurement result accurately and reliably, it is necessary to before the use, simulation is actual
Operating mode is calibrated to static probe.In defense-related science, technology and industry field, the accurate measurement of air-flow static pressure is related in Project R＆D
The acquisition of important foundation data.Therefore, it is necessary to carry out calibration to static probe under the conditions of supersonic speed, solves the measurement of static pressure
Amendment problem, its accuracy of measurement is improved, so as to obtain real supersonic flow field information.
The content of the invention
It is an object of the invention to provide a kind of experiment for carrying out calibration to static probe under supersonic airstream environment
Device, for obtaining the static pressure coefficient of static probe, test probe actually makes to the insensitivity of flow-deviation angle degree for improving
With the confidence level and measurement accuracy of middle test data.
The purpose of the present invention is achieved through the following technical solutions.
Supersonic speed exerciser, including supersonic nozzle, test section, exhaust are used in a kind of static probe calibration disclosed by the invention
Section and standard static probe.Described test section air inlet oral area is adapted with supersonic nozzle gas outlet, test section gas outlet
It is adapted with exhaust section air inlet, upper and lower both sides are respectively symmetrically provided with for standard static probe and quiet by school in the middle part of test section
Press the hole of probe insertion.It is referred to as probe for ease of description, described standard static probe and by school static probe.Supersonic nozzle
With test section fixed seal connection, test section and exhaust section fixed seal connection, standard static probe and by school static probe point
Not Cha Ru test section both sides up and down symmetrical holes, probe and test section fixing seal at patchhole.
To realize insensitivity of the measurement probe to flow-deviation angle degree, increase probe chucking device and bridge type bracket, also
Movable sealing structure need to be increased at the patchhole of test section.Described probe chucking device is used to realize that probe is loaded and angle becomes
Change, be two degrees of freedom displacement mechanism.Probe chucking device ensures the relative position with test section by bridge type bracket.Described examination
Testing section includes test section main body and movable sealing structure.Described movable sealing structure is used to realize probe rotation and sealing.
Described movable sealing structure includes clamp nut, gland, rubber seal, plastic seal ring, flange.Flange and examination
A section main body fixing seal is tested, the hole that probe is inserted on flange, plastic seal ring, rubber seal, gland, clamp nut cover successively
On probe, compressed by clamp nut.
For real-time monitoring test section axial flow velocity, static pressure survey is in axial direction equally spaced in test section main body
Hole.
When static pressure is measured in supersonic airstream, shock wave is produced in front of the probe of probe, in order to reduce error, is preferably adopted
With the probe of pointed cone form, and drift angle is limited no more than 10 °, to ensure that shock wave depends on spy in Mach number working range
Crown end, in oblique shock wave form.Air-flow is compressed when passing through shock wave to raise static pressure, and when then passing through probe shoulder, air-flow is swollen
It is swollen to reduce static pressure, the pressure tap for influenceing the place of cancelling out each other and probe being set in air-flow to static pressure.
Described supersonic nozzle is preferably as follows method realization：
To obtain supersonic airstream, using the supersonic nozzle of converging diverging, air-flow is changed into Supersonic after supersonic nozzle
Speed, the flow field of default gasflow mach number and quality requirements is met in the position of test section.
Supersonic nozzle is designed using the method for characteristic curves, and method of characteristic curves design general principle is：In supersonic flow field, Shunping County
The air-flow in face, which runs into interior curved knuckle, will produce compressional wave, run into the knuckle of excurvation and will produce dilatational wave.According to Concept of Limit, energy
It is enough that jet pipe is divided into unlimited number of short broken line, there is an infinitely small knuckle between the adjacent line segment of each two.Initial
Section, because each knuckle is bent outward, so producing unlimited faint dilatational wave at knuckle.And terminating section, each knuckle
All it is inflexed, therefore a unlimited faint compressional wave is produced at knuckle.If it is even straight ultrasound at initial segment
Air-flow, then design terminate section curve, so that dilatational wave caused by initial segment is all terminated compressional wave caused by section and offset, then
Uniform supersonic flow field is obtained in nozzle exit.But characteristic line method is not suitable for subsonic speed situation, design can only be from venturi
Sonic flow start.It is designed to obtain the supersonic nozzle of ideal Mach number using the method for characteristic curves.
Typical supersonic nozzle is divided into contraction section, initial segment and expansion segment three parts.Supersonic nozzle requires to reach larynx
The sonic flow in portion must be uniform.If stable section incoming is uniform, as long as there is smooth continuous and gradual change a receipts
Contracting curve is with regard to that can substantially meet requirement.Contraction section is designed using this pungent base curves formula of Vito, when contraction section entrance, throat
Size and shrink segment length it is given after, this pungent base curves shape of Vito i.e. it has been determined that by way of moving axle controlling curve shape
Shape.Vito Xin Siji formula are：
In formula,
R-contraction section sweep；
r1- contraction section curve starting point radius；
r0- contraction section curve end point radius；
L-contraction section length of curve；
With a distance from x-from contraction section curve starting point.
Initial segment and termination section belong to nozzle divergence cone, and it is that ABC expands molded line to define expansion segment curve, is had at B points
Maximum slope, B are referred to as turning point, and the curve before turning point B claims initial segment, and later curve is terminates section.A points are jet pipe
The starting point of expansion segment curve, C points are terminating point.
The design of initial segment curve is to the velocity of sound incoming of throat be become the source stream at turning point, and flowing source point is located at
On nozzle axis, centered on source point O ', passing through has identical Mach number on BB ' circular arcs, and airflow direction is along radial direction.
Initial segment is designed using Foelsch methods.Air-flow forms uniform source after initial segment on the BB ' lines of end
Stream, it has been experienced that, analytic curve is used between throat and turning point, allows to obtain approximate source stream in turning point.Foelsch
Method needs first to determine turning point B, lengthwise position and the maximum swelling angle beta of selectionBIt is relevant, while have with nozzle exit Mach number
Close, in Ma ＜ 5, Foelsch is preferentially used：
In formula：ν1The Prandtl-Meyer angle of-design Mach number, tables look-up；
Y*-throat radius；
H-nozzle exit radius.
Turning point B coordinate xB、yBIt is calculated according to following formula：
yE=rEsinβE
rBDetermined by source stream area governing equation, i.e.,：
In formula：σ --- for axisymmetric nozzle σ=1, two-dimensional nozzle σ=0；
MaB--- Mach number at B points；
γ --- Ratio of Specific Heats of Air, air 1.4, combustion gas 1.33.
MaBBy corresponding Prandtl-Meyer angle νBTable look-up, νBIt can be tried to achieve according to following formula：
It is as follows using experience curve equation to ensure to meet source stream condition at turning point：
Jet pipe initial segment curve has formed complete source stream designed for ensureing that air-flow reaches at turning point B, then terminates
Section curve BC design is the uniform flowing parallel to axis for described supersonic speed source stream being transformed into needs.
In order to obtain having default Mach number and parallel to the uniform flow of wind-tunnel axis in spout outlet, BC section walls must
The each dilatational wave reached thereon need be eliminated, it is not reflected.According to the concept of the limit, BC, line are divided into unlimited more
Individual unlimited short broken line, it is unlimited faint compressional wave caused by each knuckle.In supersonic flow field, all are faint to disturb
Dynamic propagated along mach line, and faint ripple is also a mach line.In supersonic flows, characteristic curve weighs everywhere with mach line
Close, therefore characteristic curve is exactly mach line.The right lateral characteristic curve sent by B points meets at E with nozzle axis, it is clear that the influence of BC section curves
Can only be in the downstream of BE lines, flowing is still source stream in BEB ' B areas, and certain BE lines are not straight lines.
In source stream area, had according to area than relation：
For the air-flow of M points, then behind β angles of transferring, it is changed into parallel to wind-tunnel axle, and reaches the uniform of test section Mach number
Air-flow.Because β is in 0≤β≤βBIn the range of change, given β can be obtained by ν, and corresponding Mach number and r be also just really on BE lines
It is fixed, thus the every bit on BE lines is all to determine.
So far, described supersonic nozzle internal face Curve Design is completed.
Beneficial effect：
1. supersonic speed exerciser is used in a kind of static probe calibration disclosed by the invention, static pressure under the conditions of supersonic speed can be realized
The calibration problem of probe, the amendment of result is measured to static probe, and then improve the accuracy of static probe measurement.
2. a kind of static probe calibration supersonic speed exerciser disclosed by the invention, increases probe chucking device and bridge type branch
Frame, increase movable sealing structure at the patchhole of test section, by the rotation of angle known to being carried out to probe, meanwhile, contrast mark
Quasi- static pressure, the unwise sensitivity to angle windward by school static probe is grasped, static probe is tested in supersonic airstream environment to meeting
The unwise sensitivity at wind angle, foundation is provided for the actual use installation of static probe.
Brief description of the drawings
Fig. 1 is schematic structural view of the invention；
Fig. 2 is supersonic nozzle expansion segment schematic diagram；
Fig. 3 is moderate supersonic speed jet pipe of the present invention, test section, exhaust section structural representation；
Fig. 4 is the movable sealing structure schematic diagram of test section in the present invention；
Fig. 5 is Plays static probe head construction schematic diagram of the present invention.
Wherein：1-supersonic nozzle, 2-test section, 2.1-test section main body, 2.1.1- static openings, 2.1.2- are visited
Pin mounting seat, 2.2-movable sealing structure, 2.2.1- clamp nuts, 2.2.2- glands, 2.2.3- rubber seal, 2.2.4- modelings
Expect sealing ring, 2.2.5- flanges, 3- exhaust sections, 4- probes chucking device, 5- bridge type brackets, 6- standards static probe, 7- by school
Static probe.
Embodiment
In order to better illustrate objects and advantages of the present invention, the content of the invention is done further with example below in conjunction with the accompanying drawings
Explanation.
Embodiment 1：
A kind of static probe calibration supersonic speed exerciser disclosed in the present embodiment, including supersonic nozzle 1, test section 2,
Exhaust section 3 and standard static probe 6.The described air inlet oral area of test section 2 is adapted with the gas outlet of supersonic nozzle 1, experiment
2 gas outlets of section are adapted with the air inlet of exhaust section 3, are respectively symmetrically provided with for standard static pressure both sides up and down at the middle part of test section 2
Probe 6 and the hole inserted by school static probe 7.For ease of description, described standard static probe 6 and united by school static probe 7
Claim probe.Supersonic nozzle 1 and the fixed seal connection of test section 2, test section 2 and the fixed seal connection of exhaust section 3, standard static pressure
Probe 6 and the symmetrical holes for being inserted test section both sides about 2 respectively by school static probe 7, probe are consolidated with test section 2 at patchhole
Fixed sealing.
The design of exerciser need to take into account aeroperformance, usability and manufacturability, its pneumatic design need to consider it is following because
Element：Reach the design Mach number of requirement；The section Mach Number Distribution of guarantee test section 2 is uniform；Exerciser outlet airflow direction with
Wind-tunnel diameter parallel and ensure suitable exerciser length etc..To ensure normally to couple with existing wind-tunnel, to test parts chi
It is very little to make limitation：
A) exerciser inlet diameter：150mm；
B) exerciser outlet diameter：75mm；
C) exerciser length：No more than 750mm.
Exerciser stage casing is test section 2, and it is the critical piece for carrying out calibration test.The porch of test section 2 is provided with 4 footpaths
To static opening 2.1.1, circumference uniform distribution；The axis direction of test section 2 is provided with 11 static opening 2.1.1, axial static opening
2.1.1 alignd with one in radial direction static opening 2.1.1.Static opening 2.1.1 requires vertical with internal face, and aperture keeps sharp
Angle, aperture 1mm.Mouth is connect in the test section 2 outside wall surface welding external screw thread corresponding with static opening 2.1.1, by pneumatic fast
Joint draws pressure.
Probe mounting seat 2.1.2 is set at the stage casing of test section 2 and axially vertical place, and aperture is Φ 30mm, outside the external welding of hole
Footpath Φ 95mm prominent face flange.
Using supersonic nozzle 1, test section 2 and exhaust section 3 as a part in the design, connected using welding procedure,
It is easily installed and processes.Exhaust section 3 is expansion shape pipeline section, is welded to connect with test section 2.The angle of flare of exhaust section 3 is designed as being less than
15 °, it is ensured that air-flow will not be produced with wall and separated.Experiment is discharged in the pipeline of wind-tunnel rear portion by exhaust section 3 with air-flow.Due to this
The gasflow mach number of test section 2 is no more than 2.0, it is possible to using the expansionary channel structure of simple diffusion.
To realize insensitivity of the measurement probe to flow-deviation angle degree, increase probe chucking device 4 and bridge type bracket 5,
Movable sealing structure 2.2 need to also be increased at the patchhole of test section 2.Described probe chucking device 4 is used to realize that probe is loaded
And angular transformation, it is two degrees of freedom displacement mechanism.Probe chucking device 4 ensures the relative position with test section 2 by bridge type bracket 5
Put.The described test section 2 of connection includes test section main body 2.1 and movable sealing structure 2.2.Described movable sealing structure 2.2 is used for
Realize probe rotation and sealing.
Probe chucking device 4 is for realizing that flow field moves radially measurement and is calibrated the portion of probe angulation deflection
Part.The probe structure that is loaded uses screw-type clasping structure, internal diameter 4.2mm, slightly larger than the pressure guiding pipe of standard static probe 6.Vertically
Displacement mechanism is driven using motor, effective travel 200mm, precision 0.01mm.Driven around Y-axis rotating mechanism using motor,
Effective angle is ± 30 °, and precision is 0.05 °.Motion controller is used to control vertical displacement mechanism and around Y-axis rotating mechanism, adopts
Two kinds of control models are controlled with field control and teletransmission.
Probe chucking device 4 is arranged on the top of static probe mounting seat by bridge type bracket 5, and bridge type bracket 5 passes through spiral shell
Nail is fixed with flange 2.2.5 before and after exerciser, and during installation, the probe of probe chucking device 4 structure that is loaded is concentric with static probe.
It is close that described movable sealing structure 2.2 includes clamp nut 2.2.1, gland 2.2.2, rubber seal 2.2.3, plastics
Seal 2.2.4, flange 2.2.5.Flange 2.2.5 and the fixing seal of test section main body 2.1, the hole that probe is inserted on flange 2.2.5,
Plastic seal ring 2.2.4, rubber seal 2.2.3, gland 2.2.2, clamp nut 2.2.1 are sequentially sleeved on probe, pass through pressure
Tight nut 2.2.1 is compressed.In actual use, the structure of viton seal ring is matched somebody with somebody using teflin ring, with screw thread pressure
Tight mode connects, and realizes dynamic sealing requirement.The structure provides lubrication using teflin ring, due to polytetrafluoroethylene (PTFE)
Sealing ring is elastic poor in itself, is also easy to produce larger residual deformation, therefore added rubber packing ring, outer rubber are employed in structure
The effect of packing ring is the elasticity for subsidizing teflin ring, to ensure that it compresses to axle surface elasticity and in matrix
The air-tightness on surface and teflin ring junction.When in use, can only be by increasing the compression on sealing contact surface
Power ensures to seal, and the thrust of sealing ring is provided by clamp nut 2.2.1, so as to ensureing the dynamic sealing of pressure guiding pipe.
For the real-time axial flow velocity of monitoring test section 2, in axial direction it is equally spaced in test section main body 2.1 quiet
Press gaging hole 2.1.1.
Described supersonic nozzle 1 is preferably as follows method realization：
To obtain supersonic airstream, using the supersonic nozzle 1 of converging diverging, air-flow is changed into super after supersonic nozzle 1
Velocity of sound, the flow field of default gasflow mach number and quality requirements is met in the position of test section 2.
Supersonic nozzle 1 is designed using the method for characteristic curves, and method of characteristic curves design general principle is：It is suitable in supersonic flow field
The air-flow of plane, which runs into interior curved knuckle, will produce compressional wave, run into the knuckle of excurvation and will produce dilatational wave.According to Concept of Limit,
Jet pipe can be divided into unlimited number of short broken line, there is an infinitely small knuckle between the adjacent line segment of each two.First
Beginning section, because each knuckle is bent outward, so producing unlimited faint dilatational wave at knuckle.And terminating section, each folding
Angle is all inflexed, therefore a unlimited faint compressional wave is produced at knuckle.If it is even straight surpass at initial segment
Information stream, then design terminate section curve, so that dilatational wave caused by initial segment is all terminated compressional wave caused by section and offset,
Then uniform supersonic flow field is obtained in nozzle exit.But characteristic line method is not suitable for subsonic speed situation, design can only be from larynx
The sonic flow in road starts.It is designed to obtain the supersonic nozzle 1 of ideal Mach number using the method for characteristic curves.
Typical supersonic nozzle 1 divides for contraction section, initial segment and expansion segment three parts.Supersonic nozzle 1 requires to reach
The sonic flow of throat must be uniform.If stable section incoming is uniform, if having one it is smooth continuous and gradual change
Shrinkage curve is with regard to that can substantially meet requirement.Contraction section is designed using this pungent base curves formula of Vito, when contraction section entrance, larynx
Portion's size and shrink segment length it is given after, this pungent base curves shape of Vito i.e. it has been determined that by way of moving axle controlling curve
Shape.Vito Xin Siji formula are：
In formula,
R-contraction section sweep；
r1- contraction section curve starting point radius；
r0- contraction section curve end point radius；
L-contraction section length of curve；
With a distance from x-from contraction section curve starting point.
Initial segment and termination section belong to nozzle divergence cone, and it is that ABC expands molded line to define expansion segment curve, is had at B points
Maximum slope, B are referred to as turning point, and the curve before turning point B claims initial segment, and later curve is terminates section.A points are jet pipe
The starting point of expansion segment curve, C points are terminating point.
The design of initial segment curve is to the velocity of sound incoming of throat be become the source stream at turning point, and flowing source point is located at
On nozzle axis, centered on source point O ', passing through has identical Mach number on BB ' circular arcs, and airflow direction is along radial direction.
Initial segment is designed using Foelsch methods.Air-flow forms uniform source after initial segment on the BB ' lines of end
Stream, it has been experienced that, analytic curve is used between throat and turning point, allows to obtain approximate source stream in turning point.Foelsch
Method needs first to determine turning point B, lengthwise position and the maximum swelling angle beta of selectionBIt is relevant, while have with nozzle exit Mach number
Close, in Ma ＜ 5, Foelsch is preferentially used：
In formula：ν1The Prandtl-Meyer angle of-design Mach number, tables look-up；
y*- throat radius；
H-nozzle exit radius.
Turning point B coordinate xB、yBIt is calculated according to following formula：
yE=rEsinβE
rBDetermined by source stream area governing equation, i.e.,：
In formula：σ --- for axisymmetric nozzle σ=1, two-dimensional nozzle σ=0；
MaB--- Mach number at B points；
γ --- Ratio of Specific Heats of Air, air 1.4, combustion gas 1.33.
MaBCan be by corresponding Prandtl-Meyer angle νBTable look-up, νBIt can be tried to achieve according to following formula：
It is as follows using experience curve equation to ensure to meet source stream condition at turning point：
Jet pipe initial segment curve has formed complete source stream designed for ensureing that air-flow reaches at turning point B, then terminates
Section curve BC design is the uniform flowing parallel to axis for described supersonic speed source stream being transformed into needs.
In order to obtain having certain Mach number and parallel to the uniform flow of wind-tunnel axis in spout outlet, BC section walls must
The each dilatational wave reached thereon need be eliminated, it is not reflected.According to the concept of the limit, BC, line are divided into unlimited more
Individual unlimited short broken line, it is unlimited faint compressional wave caused by each knuckle.In supersonic flow field, all are faint to disturb
Dynamic propagated along mach line, and faint ripple is also a mach line.In supersonic flows, characteristic curve weighs everywhere with mach line
Close, therefore characteristic curve is exactly mach line.The right lateral characteristic curve sent by B points meets at E with nozzle axis, it is clear that the influence of BC section curves
Can only be in the downstream of BE lines, flowing is still source stream in BEB ' B areas, and certain BE lines are not straight lines.
In source stream area, had according to area than relation：
For the air-flow of M points, then behind β angles of transferring, it is changed into parallel to wind-tunnel axle, and reaches the equal of the Mach number of test section 2
Even air-flow.Because β is in 0≤β≤βBIn the range of change, given β can be obtained by ν, and corresponding Mach number and r be also on BE lines
Determine, thus every bit on BE lines is all to determine.
So far, the described internal face Curve Design of supersonic nozzle 1 is completed.
When static pressure is measured in supersonic airstream, shock wave is produced in front of the probe of probe, in order to reduce error, is preferably adopted
With the probe of pointed cone form, and drift angle is limited no more than 10 °, to ensure that shock wave depends on spy in Mach number working range
Crown end, in oblique shock wave form.Air-flow is compressed when passing through shock wave to raise static pressure, with it is laggard cross probe shoulder when, air-flow is swollen
It is swollen to reduce static pressure, the pressure tap for influenceing the place of cancelling out each other and probe being set in air-flow to static pressure.
Standard static probe 6, using circular cone type structure, is completed according to probe design requirement through Precision Machining.Probe
It is connected by flange 2.2.5 with exerciser, flange 2.2.5 is sealed with exerciser using plain washer.To meet that probe can work
Shi Jinhang twist motions, while can also seal, probe uses movable sealing structure 2.2 with flange 2.2.5 junctions.
Above-described specific descriptions, the purpose, technical scheme and beneficial effect of invention are carried out further specifically
It is bright, it should be understood that the specific embodiment that the foregoing is only the present invention, the protection model being not intended to limit the present invention
Enclose, within the spirit and principles of the invention, any modification, equivalent substitution and improvements done etc., should be included in the present invention
Protection domain within.
Claims (6)
1. supersonic speed exerciser is used in a kind of static probe calibration, it is characterised in that：Including supersonic nozzle (1), test section (2),
Exhaust section (3) and standard static probe (6)；Described test section (2) air inlet oral area is mutually fitted with supersonic nozzle (1) gas outlet
Match somebody with somebody, test section (2) gas outlet is adapted with exhaust section (3) air inlet, and both sides are respectively symmetrically provided with up and down in the middle part of test section (2)
The hole inserted for standard static probe (6) and by school static probe (7)；For ease of description, described standard static probe (6)
It is referred to as probe with by school static probe (7)；Supersonic nozzle (1) and test section (2) fixed seal connection, test section (2) and row
Gas section (3) fixed seal connection, standard static probe (6) and is inserted test section (2) both sides up and down by school static probe (7) respectively
Symmetrical holes, probe and test section (2) fixing seal at patchhole.
2. supersonic speed exerciser is used in a kind of static probe calibration as claimed in claim 1, it is characterised in that：To realize that measurement is visited
For the insensitivity of flow-deviation angle degree, increase probe chucking device (4) and bridge type bracket (5) need to also be in test sections (2)
Increase movable sealing structure at patchhole；Described probe chucking device (4) is used to realize that probe is loaded and angular transformation, be two from
By degree displacement mechanism；Probe chucking device (4) ensures the relative position with test section (2) by bridge type bracket (5)；Described examination
Testing section (2) includes test section main body (2.1) and movable sealing structure (2.2)；Described movable sealing structure (2.2) is used to realize probe
Rotation and sealing.
3. supersonic speed exerciser is used in a kind of static probe calibration as claimed in claim 2, it is characterised in that：Described dynamic sealing
Structure (2.2) include clamp nut (2.2.1), gland (2.2.3), rubber seal (2.2.3), plastic seal ring (2.2.4),
Flange (2.2.5)；Flange (2.2.5) and test section main body (2.1) fixing seal, the hole that probe is inserted on flange (2.2.5), modeling
Material sealing ring (2.2.4), rubber seal (2.2.3), gland (2.2.3), clamp nut (2.2.1) are sequentially sleeved on probe,
Compressed by clamp nut (2.2.1).
4. supersonic speed exerciser is used in a kind of static probe calibration as claimed in claim 3, it is characterised in that：For monitoring examination in real time
Section (2) axial flow velocity is tested, static opening (2.1.1) is in axial direction equally spaced in test section main body (2.1).
5. supersonic speed exerciser is used in a kind of static probe calibration as claimed in claim 4, it is characterised in that：In supersonic airstream
During middle measurement static pressure, shock wave is produced in front of the probe of probe, in order to reduce error, using the probe of pointed cone form, and is limited
Drift angle processed is no more than 10 °, to ensure that shock wave depends on probe tip in Mach number working range, in oblique shock wave form；Gas
Stream is compressed when passing through shock wave to raise static pressure, and when then passing through probe shoulder, flow expansion reduces static pressure, in air-flow to quiet
The pressure tap for influenceing the place of cancelling out each other and probe being set of pressure.
6. supersonic speed exerciser is used in a kind of static probe calibration as described in claim 1,2,3,4 or 5, it is characterised in that：Institute
The supersonic nozzle (1) stated selects following method to realize：
To obtain supersonic airstream, using the supersonic nozzle (1) of converging diverging, air-flow is changed into super after supersonic nozzle (1)
Velocity of sound, the flow field of default gasflow mach number and quality requirements is met in the position of test section (2)；
Supersonic nozzle (1) is designed using the method for characteristic curves, and method of characteristic curves design general principle is：In supersonic flow field, Shunping County
The air-flow in face, which runs into interior curved knuckle, will produce compressional wave, run into the knuckle of excurvation and will produce dilatational wave；According to Concept of Limit, energy
It is enough that jet pipe is divided into unlimited number of short broken line, there is an infinitely small knuckle between the adjacent line segment of each two；Initial
Section, because each knuckle is bent outward, so producing unlimited faint dilatational wave at knuckle；And terminating section, each knuckle
All it is inflexed, therefore a unlimited faint compressional wave is produced at knuckle；If it is even straight ultrasound at initial segment
Air-flow, then design terminate section curve, so that dilatational wave caused by initial segment is all terminated compressional wave caused by section and offset, then
Uniform supersonic flow field is obtained in nozzle exit；But characteristic line method is not suitable for subsonic speed situation, design can only be from venturi
Sonic flow start；It is designed to obtain the supersonic nozzle (1) of ideal Mach number using the method for characteristic curves；
Typical supersonic nozzle (1) is divided into contraction section, initial segment and expansion segment three parts；Supersonic nozzle (1) requires to reach
The sonic flow of throat must be uniform；If stable section incoming is uniform, if having one it is smooth continuous and gradual change
Shrinkage curve is with regard to that can substantially meet requirement；Contraction section is designed using this pungent base curves formula of Vito, when contraction section entrance, larynx
Portion's size and shrink segment length it is given after, this pungent base curves shape of Vito i.e. it has been determined that by way of moving axle controlling curve
Shape；Vito Xin Siji formula are：
<mrow>
<mi>r</mi>
<mo>=</mo>
<mfrac>
<msub>
<mi>r</mi>
<mn>0</mn>
</msub>
<msqrt>
<mrow>
<mn>1</mn>
<mo>-</mo>
<mo>&lsqb;</mo>
<mn>1</mn>
<mo>-</mo>
<msup>
<mrow>
<mo>(</mo>
<mfrac>
<msub>
<mi>r</mi>
<mn>0</mn>
</msub>
<msub>
<mi>r</mi>
<mn>1</mn>
</msub>
</mfrac>
<mo>)</mo>
</mrow>
<mn>2</mn>
</msup>
<mo>&rsqb;</mo>
<mfrac>
<msup>
<mrow>
<mo>(</mo>
<mrow>
<mn>1</mn>
<mo>-</mo>
<mn>3</mn>
<msup>
<mi>x</mi>
<mn>2</mn>
</msup>
<mo>/</mo>
<msup>
<mi>a</mi>
<mn>2</mn>
</msup>
</mrow>
<mo>)</mo>
</mrow>
<mn>2</mn>
</msup>
<msup>
<mrow>
<mo>(</mo>
<mrow>
<mn>1</mn>
<mo>+</mo>
<msup>
<mi>x</mi>
<mn>2</mn>
</msup>
<mo>/</mo>
<msup>
<mi>a</mi>
<mn>2</mn>
</msup>
</mrow>
<mo>)</mo>
</mrow>
<mn>3</mn>
</msup>
</mfrac>
</mrow>
</msqrt>
</mfrac>
</mrow>
In formula,
R-contraction section sweep；
r1- contraction section curve starting point radius；
r0- contraction section curve end point radius；
<mrow>
<mi>a</mi>
<mo>=</mo>
<msqrt>
<mn>3</mn>
</msqrt>
<mi>l</mi>
<mo>;</mo>
</mrow>
L-contraction section length of curve；
With a distance from x-from contraction section curve starting point；
Initial segment and termination section belong to nozzle divergence cone, and it is that ABC expands molded line to define expansion segment curve, has maximum at B points
Slope, B is referred to as turning point, and the curve before turning point B claims initial segment, and later curve is terminates section；A points are nozzle-divergence
The starting point of section curve, C points are terminating point；
The design of initial segment curve is to the velocity of sound incoming of throat be become the source stream at turning point, and flowing source point is located at jet pipe
On axis, centered on source point O ', passing through has identical Mach number on BB ' circular arcs, and airflow direction is along radial direction；
Initial segment is designed using Foelsch methods；Air-flow forms uniform source stream after initial segment on the BB ' lines of end,
Experience have shown that using analytic curve between throat and turning point, allow to obtain approximate source stream in turning point；Foelsch side
Method needs first to determine turning point B, lengthwise position and the maximum swelling angle beta of selectionBIt is relevant, while have with nozzle exit Mach number
Close, in Ma ＜ 5, Foelsch is used：
<mrow>
<msub>
<mi>&beta;</mi>
<mi>B</mi>
</msub>
<mo>=</mo>
<mfrac>
<mn>1</mn>
<mn>2</mn>
</mfrac>
<msub>
<mi>v</mi>
<mn>1</mn>
</msub>
<msup>
<mrow>
<mo>(</mo>
<mfrac>
<msup>
<mi>y</mi>
<mo>*</mo>
</msup>
<mi>h</mi>
</mfrac>
<mo>)</mo>
</mrow>
<mfrac>
<mn>2</mn>
<mn>9</mn>
</mfrac>
</msup>
</mrow>
In formula：ν1The Prandtl-Meyer angle of-design Mach number, tables look-up；
Y*-throat radius；
H-nozzle exit radius；
Turning point B coordinate xB、yBIt is calculated according to following formula：
<mrow>
<msub>
<mi>x</mi>
<mi>B</mi>
</msub>
<mo>=</mo>
<mfrac>
<mrow>
<mn>3</mn>
<mrow>
<mo>(</mo>
<msub>
<mi>y</mi>
<mi>B</mi>
</msub>
<mo>-</mo>
<msup>
<mi>y</mi>
<mo>*</mo>
</msup>
<mo>)</mo>
</mrow>
</mrow>
<mrow>
<mn>2</mn>
<msub>
<mi>tan&beta;</mi>
<mi>B</mi>
</msub>
</mrow>
</mfrac>
</mrow>
yB=rBsinβB
rBDetermined by source stream area governing equation, i.e.,：
<mrow>
<msup>
<mrow>
<mo>(</mo>
<mfrac>
<msub>
<mi>r</mi>
<mi>B</mi>
</msub>
<msub>
<mi>r</mi>
<mn>0</mn>
</msub>
</mfrac>
<mo>)</mo>
</mrow>
<mrow>
<mn>1</mn>
<mo>+</mo>
<mi>&sigma;</mi>
</mrow>
</msup>
<mo>=</mo>
<mfrac>
<mn>1</mn>
<mrow>
<msub>
<mi>Ma</mi>
<mi>B</mi>
</msub>
</mrow>
</mfrac>
<msup>
<mrow>
<mo>&lsqb;</mo>
<mrow>
<mo>(</mo>
<mfrac>
<mn>2</mn>
<mrow>
<mi>&gamma;</mi>
<mo>+</mo>
<mn>1</mn>
</mrow>
</mfrac>
<mo>)</mo>
</mrow>
<mrow>
<mo>(</mo>
<mn>1</mn>
<mo>+</mo>
<mfrac>
<mrow>
<mi>&gamma;</mi>
<mo>-</mo>
<mn>1</mn>
</mrow>
<mn>2</mn>
</mfrac>
<msup>
<msub>
<mi>Ma</mi>
<mi>B</mi>
</msub>
<mn>2</mn>
</msup>
<mo>)</mo>
</mrow>
<mo>&rsqb;</mo>
</mrow>
<mfrac>
<mrow>
<mi>&gamma;</mi>
<mo>+</mo>
<mn>1</mn>
</mrow>
<mrow>
<mn>2</mn>
<mrow>
<mo>(</mo>
<mrow>
<mi>&gamma;</mi>
<mo>-</mo>
<mn>1</mn>
</mrow>
<mo>)</mo>
</mrow>
</mrow>
</mfrac>
</msup>
</mrow>
In formula：σ --- for axisymmetric nozzle σ=1, two-dimensional nozzle σ=0；
MaB--- Mach number at B points；
γ --- Ratio of Specific Heats of Air, air 1.4, combustion gas 1.33；
MaBBy corresponding Prandtl-Meyer angle νBTable look-up, νBTried to achieve according to following formula：
<mrow>
<msub>
<mi>&upsi;</mi>
<mi>B</mi>
</msub>
<mo>=</mo>
<mfrac>
<mrow>
<msub>
<mi>&upsi;</mi>
<mn>1</mn>
</msub>
<mo>-</mo>
<msub>
<mi>&beta;</mi>
<mi>B</mi>
</msub>
</mrow>
<mrow>
<mn>1</mn>
<mo>+</mo>
<mi>&sigma;</mi>
</mrow>
</mfrac>
</mrow>
It is as follows using experience curve equation to ensure to meet source stream condition at turning point：
<mrow>
<mi>y</mi>
<mo>=</mo>
<msup>
<mi>y</mi>
<mo>*</mo>
</msup>
<mo>+</mo>
<mrow>
<mo>(</mo>
<mfrac>
<mrow>
<msub>
<mi>tan&beta;</mi>
<mi>B</mi>
</msub>
</mrow>
<msub>
<mi>x</mi>
<mi>B</mi>
</msub>
</mfrac>
<mo>)</mo>
</mrow>
<msup>
<mi>x</mi>
<mn>2</mn>
</msup>
<mrow>
<mo>(</mo>
<mrow>
<mn>1</mn>
<mo>-</mo>
<mfrac>
<mi>x</mi>
<mrow>
<mn>3</mn>
<msub>
<mi>x</mi>
<mi>B</mi>
</msub>
</mrow>
</mfrac>
</mrow>
<mo>)</mo>
</mrow>
</mrow>
Jet pipe initial segment curve has formed complete source stream designed for ensureing that air-flow reaches at turning point B, then terminates Duan Qu
Line BC design is the uniform flowing parallel to axis for described supersonic speed source stream being transformed into needs；
In order in spout outlet obtain that there is default Mach number and be had to parallel to the uniform flow of wind-tunnel axis, BC section walls
The each dilatational wave reached thereon is eliminated, it is not reflected；According to the concept of the limit, BC, line are divided into unlimited number of nothing
Short broken line is limited, is unlimited faint compressional wave caused by each knuckle；In supersonic flow field, all faint disturbances are all
It is to be propagated along mach line, faint ripple is also a mach line；In supersonic flows, characteristic curve overlaps everywhere with mach line,
Therefore characteristic curve is exactly mach line；The right lateral characteristic curve sent by B points meets at E with nozzle axis, it is clear that the influence of BC section curves is only
Can be in the downstream of BE lines, flowing is still source stream in BEB ' B areas, and certain BE lines are not straight lines；
In source stream area, had according to area than relation：
<mrow>
<msup>
<mrow>
<mo>(</mo>
<mfrac>
<mi>r</mi>
<msub>
<mi>r</mi>
<mn>0</mn>
</msub>
</mfrac>
<mo>)</mo>
</mrow>
<mrow>
<mn>1</mn>
<mo>+</mo>
<mi>&sigma;</mi>
</mrow>
</msup>
<mo>=</mo>
<mfrac>
<mn>1</mn>
<mrow>
<mi>M</mi>
<mi>a</mi>
</mrow>
</mfrac>
<msup>
<mrow>
<mo>(</mo>
<mrow>
<mfrac>
<mrow>
<mi>&gamma;</mi>
<mo>-</mo>
<mn>1</mn>
</mrow>
<mrow>
<mi>&gamma;</mi>
<mo>+</mo>
<mn>1</mn>
</mrow>
</mfrac>
<msup>
<mi>Ma</mi>
<mn>2</mn>
</msup>
<mo>+</mo>
<mfrac>
<mn>2</mn>
<mrow>
<mi>&gamma;</mi>
<mo>+</mo>
<mn>1</mn>
</mrow>
</mfrac>
</mrow>
<mo>)</mo>
</mrow>
<mfrac>
<mrow>
<mi>&gamma;</mi>
<mo>+</mo>
<mn>1</mn>
</mrow>
<mrow>
<mn>2</mn>
<mrow>
<mo>(</mo>
<mrow>
<mi>&gamma;</mi>
<mo>-</mo>
<mn>1</mn>
</mrow>
<mo>)</mo>
</mrow>
</mrow>
</mfrac>
</msup>
</mrow>
For the air-flow of M points, then behind β angles of transferring, it is changed into parallel to wind-tunnel axle, and reaches the uniform of test section (2) Mach number
Air-flow；Because β is in 0≤β≤βBIn the range of change, given β can be obtained by ν, and corresponding Mach number and r be also just really on BE lines
It is fixed, thus the every bit on BE lines is all to determine；
So far, described supersonic nozzle (1) internal face Curve Design is completed.
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| CN108709712A (en) * | 2018-07-31 | 2018-10-26 | 大连凌海华威科技服务有限责任公司 | Subsonic jets formula air feeders calibration wind tunnel |
| CN111077345A (en) * | 2019-12-20 | 2020-04-28 | 西安航天动力研究所 | Mach number calibration method under high-temperature supersonic velocity pure gas flow field environment |
| CN112747886A (en) * | 2020-12-29 | 2021-05-04 | 中国航天空气动力技术研究院 | Thin-wall throat |
| CN115901074A (en) * | 2022-12-13 | 2023-04-04 | 重庆大学 | Movable probe device for measuring pressure in flow channel of spray pipe |
| CN116256143A (en) * | 2023-05-15 | 2023-06-13 | 中国航空工业集团公司沈阳空气动力研究所 | Integrated structure of pipe wind tunnel spray pipe and test section and operation method |
| CN117949138A (en) * | 2024-03-27 | 2024-04-30 | 中国航空工业集团公司沈阳空气动力研究所 | System and method for calibrating dynamic pressure of pipe wind tunnel in high-temperature environment |
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| CN108709712A (en) * | 2018-07-31 | 2018-10-26 | 大连凌海华威科技服务有限责任公司 | Subsonic jets formula air feeders calibration wind tunnel |
| CN111077345A (en) * | 2019-12-20 | 2020-04-28 | 西安航天动力研究所 | Mach number calibration method under high-temperature supersonic velocity pure gas flow field environment |
| CN112747886A (en) * | 2020-12-29 | 2021-05-04 | 中国航天空气动力技术研究院 | Thin-wall throat |
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| CN115901074A (en) * | 2022-12-13 | 2023-04-04 | 重庆大学 | Movable probe device for measuring pressure in flow channel of spray pipe |
| CN115901074B (en) * | 2022-12-13 | 2024-06-04 | 重庆大学 | Movable probe device for measuring pressure in spray pipe flow channel |
| CN116256143A (en) * | 2023-05-15 | 2023-06-13 | 中国航空工业集团公司沈阳空气动力研究所 | Integrated structure of pipe wind tunnel spray pipe and test section and operation method |
| CN116256143B (en) * | 2023-05-15 | 2023-07-14 | 中国航空工业集团公司沈阳空气动力研究所 | Integrated structure of pipe wind tunnel spray pipe and test section and operation method |
| CN117949138A (en) * | 2024-03-27 | 2024-04-30 | 中国航空工业集团公司沈阳空气动力研究所 | System and method for calibrating dynamic pressure of pipe wind tunnel in high-temperature environment |
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